Safety is probably the most important factor in the design and operation of aeronautical systems. Mechanical failure of an airplane component can result in a catastrophic demise of the entire vessel, frequently causing huge loss of life, and necessitates application of significant resources to analyze the causes of the failure so that recurrences in other systems can be mitigated. The economic loss propagates far beyond the immediate loss of the afflicted airplane and post-failure investigation, however. After the failure of a particular component in a particular airplane, industry regulations will usually mandate withdrawal from service of all instances of the same component that have seen at least the same number of hours of operation in other airplanes as had the actual component that failed. Overall safety considerations may also mandate that particular components are retired from service once they have reached a certain number of hours of operation, even though that number of hours may be considerably lower than the typical period when failure might be expected—often by as wide a margin as a factor of three. Nevertheless, reacting to the failure, or forced retirement, of a type of component at the industry-wide scale is time-consuming and expensive, but may also not be strictly necessary because many of the components are replaced when they still have many hundreds or thousands of hours of effective and reliable service life left. Thus, alternative approaches to management of component life are called for.
The reliability of jet engine components has a special status because of the extreme conditions of operation of the components. Fatigue-related failure of a jet engine component such as turbine blade can arise from several mechanisms. For a discussion of modeling fatigue crack propagation in an aircraft engine fan blade attachment, see for example Barlow and Chandra, Int. J. Fatigue, 27:1661-1668 (2005). Just as with many other high speed rotating systems, current safety regulations require gas turbine engines to satisfy both crack initiation (safe-life) and fatigue crack growth (damage tolerant) design criteria. In an attempt to satisfy these requirements, non-destructive inspection techniques such as fluorescent penetrant, eddy current, and ultrasonic inspections have been implemented by aviation organizations such as the United States Air Force (USAF), to detect small cracks at critical locations. These approaches, which rely on systematic inspections of critical life-limiting locations in components, detect cracks that can potentially grow to failure within the next inspection interval. However, such non-destructive inspections cannot be performed in-service and require a complete disassembly of the engine, which is extremely time-consuming and therefore also expensive. Hence, the current approach to life management of components in service is both time consuming and expensive.
Component reliability has also been tested during the design phase. Conventional durability testing of engine components is carried out with full-scale mock-ups in dedicated facilities and is also an extremely costly and time-consuming affair—typically costing several million dollars and requiring a year or more to complete. Hitherto it has been believed that only a full-scale test rig can accurately replicate the thermal gradients and stresses that are actually placed on a turbine component when in service. Testing a component can therefore only be carried out with a very small number of copies—usually as few as one or two—of the component. Testing to failure is needed to fully assess the durability of components, however, but is seldom if ever performed because of the damage that is wrought on the whole engine and the test-environment during just a single failure. Approval of a final design, however, requires at least one (and preferably many) non-failure durability tests on the near-final state of the turbine engine design. If durability issues are identified at a late stage in the design process, redesigns are expensive and have tremendous impact on delivery schedule to aircraft manufacturers and subsequent aircraft purchasers.
Failure of a component can arise from a combination of causes related to geometry, thermal conditions, and constituent materials. Thus, testing of the materials themselves also plays a vital role in overall reliability assessment. Small coupon tests—tests on standardized representative samples of material—have traditionally been used to determine damage mechanisms. The geometry of the coupons is usually simple, however, and typically comprises shapes such as smooth round bars, or flat bars, both with or without simple notches. Simple coupons cannot properly simulate the complex geometry of actual hardware. Furthermore, the coupons are typically very small in size (e.g., a volume less than ¼ cubic inch, such as ⅛ cubic inch) so that they can be heated and cooled rapidly over many thousands of cycles, although the tests are usually performed at a constant temperature using laboratory air or a vacuum. During testing, the coupons are usually subjected only to simple loading such as push-pull testing or rotating-bending in which the loading is cycled between constant values of the minimum and maximum stress. Coupon tests are valuable in determining potential damage mechanisms, and evaluating the relative resistance of different materials, but because the coupons do not have the structural geometry, loading conditions, and environment of actual components, they cannot be relied upon to provide detailed predictions of structural durability of those actual components.
Because there is so much difference between a coupon test and a full-scale engine durability test, an intermediate level of testing is often performed. One common intermediate test is the spin test. This test uses individual full-size components (rather than a complete engine). A component tested in a spin test has a geometry identical or similar to the actual components in operational service, so that loads resulting from shape and size are accurately represented. These tests are typically performed in a vacuum at constant temperatures. Thus, loading from thermal gradients and other environmental influences are not represented. Even an intermediate test that involves just spinning hardware in a constant temperature vacuum pit can cost several hundred thousand dollars and up to a year to complete. These costs and schedule requirements have a drastic influence on the amount and extent of durability testing that can be performed.
Accordingly, there is a need for a testing regime for jet engine components that is inexpensive, affords the possibility of obtaining a statistically significant number of datapoints, that can be easily performed early in the design phase of an engine, and that can facilitate reliable prediction of engine life.
The discussion of the background to the invention herein is included to explain the context of the invention. This is not to be taken as an admission that any of the material referred to was published, known, or part of the common general knowledge as at the priority date of any of the claims.
Throughout the description and claims of the specification the word “comprise” and variations thereof, such as “comprising” and “comprises”, is not intended to exclude other additives, components, integers or steps.